Shielded anti-icing system and methods
An anti-icing system is provided for an inlet lip annularly extending about a nacelle of an aircraft engine assembly. The anti-icing system includes an interior wall structure at least partially forming an annular anti-icing chamber with the inlet lip and an annular shield with a first end coupled to the interior wall structure and a second end extending into the annular anti-icing chamber. The annular shield divides the annular anti-icing chamber into first and second chamber portions fluidly coupled together by a passage formed between the second end and the internal surface of the inlet lip. The anti-icing system further includes a nozzle at an inwardly radial position relative to the first end of the annular shield such that the heated air exits the nozzle into the first chamber portion in which the annular shield blocks direct impingement on the internal surface of the inlet lip.
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The present invention generally relates to aircraft systems and methods, and more particularly relates to anti-icing systems and methods for aircraft engine assemblies.
BACKGROUNDAircraft operate in many different types of weather and conditions, including icing conditions. Ice that forms on the leading edge of a nacelle that houses an aircraft engine may break off and enter the engine, potentially resulting in undesirable issues. In addition it may adversely affect air flow into the engine, potentially resulting in reduced performance and/or negatively impacting operating margins of the aircraft. Anti-icing systems have been developed to address ice formation. Although these conventional anti-icing systems have been effective, challenges remain, particularly with respect to air flow efficiency, asymmetric or uneven heating, localized hot spots, weight, complexity, and cost.
Accordingly, it is desirable to provide improved anti-icing systems and methods. Furthermore, other desirable features and characteristics of the present invention will become apparent from the subsequent detailed description of the invention and the appended claims, taken in conjunction with the accompanying drawings and this background of the invention.
BRIEF SUMMARYIn accordance with an exemplary embodiment, an anti-icing system is provided for an inlet lip annularly extending about a nacelle of an aircraft engine assembly. The anti-icing system includes an interior wall structure at least partially forming an annular anti-icing chamber with an internal surface of the inlet lip and an annular shield with a first end coupled to the interior wall structure and a second end extending into the annular anti-icing chamber. The annular shield divides the annular anti-icing chamber into a first chamber portion and a second chamber portion such that the first and second chamber portions are fluidly coupled together by a passage formed between the second end and the internal surface of the inlet lip. The anti-icing system further includes a nozzle configured to be coupled to a source of heated air and extending through the interior wall at an inwardly radial position relative to the first end of the annular shield such that the heated air exits the nozzle into the first chamber portion in which the annular shield blocks direct impingement by the heated air on the internal surface of the inlet lip.
In accordance with an exemplary embodiment, an engine assembly includes an engine; a nacelle housing the engine and having a leading portion formed by an inlet lip; and an anti-icing system configured to direct air from the engine to the inlet lip for anti-icing the inlet lip. The anti-icing system includes an interior wall structure at least partially forming an annular anti-icing chamber with an internal surface of the inlet lip; an annular shield with a first end coupled to the interior wall structure and a second end extending into the annular anti-icing chamber, the annular shield dividing the annular anti-icing chamber into a first chamber portion and a second chamber portion, the first and second chamber portions being fluidly coupled together by a passage formed between the second end and the internal surface of the inlet lip; and a nozzle configured to be coupled to a source of heated air and extending through the interior wall at an inwardly radial position relative to the first end of the annular shield such that the heated air exits the nozzle into the first chamber portion in which the annular shield blocks direct impingement by the heated air on the internal surface of the inlet lip.
The present invention will hereinafter be described in conjunction with the following drawing figures, wherein like numerals denote like elements, and
The following detailed description is merely exemplary in nature and is not intended to limit the invention or the application and uses of the invention. As used herein, the word “exemplary” means “serving as an example, instance, or illustration.” Thus, any embodiment described herein as “exemplary” is not necessarily to be construed as preferred or advantageous over other embodiments. All of the embodiments described herein are exemplary embodiments provided to enable persons skilled in the art to make or use the invention and not to limit the scope of the invention which is defined by the claims. Furthermore, there is no intention to be bound by any expressed or implied theory presented in the preceding technical field, background, brief summary, or the following detailed description.
Broadly, exemplary embodiments disclosed herein provide anti-icing systems and methods in which heated air flow is delivered by a nozzle into a chamber formed within an inlet lip of an engine assembly. In particular, exemplary embodiments include a shield within the chamber that functions to prevent the heated air exiting the nozzle from directly impinging onto the inner surface of the inlet lip. In effect, the shield divides the chamber into first and second portions to ensure proper mixing and temperature uniformity in heating the inlet lip.
The aircraft engine assembly 100 includes an engine 104 (schematically shown) housed in a nacelle 106. The nacelle 106 has an annular shape and may be formed by a number of sections or components, including a cowl section 110, an inlet cowl section 120, and an exhaust cowl section 130. The inlet cowl section 120 may be formed by an inner section 122, an outer section 124, and an inlet lip 126. The outer section 124 circumscribes the inner section 122, and the inlet lip 126 joins the inner section 122 to the outer section 124 to form the leading edge of the aircraft engine assembly 100. Generally, the inlet cowl section 120 defines an inlet opening for directing air into the engine 104. At least a portion of the air entering the assembly 100 is combusted by the engine 104 to convert a portion of the energy from the fuel to rotational energy, causing the turbine to rotate, which forces air through the engine. The resulting products are forcibly exhausted through an exhaust opening(s) at least partially defined by the exhaust cowl section 130. The force associated with this exhausted gas is called thrust.
As shown, the inlet lip 126 is the leading edge of the engine assembly 100. During some weather or operating conditions, ice has the tendency to form and accumulate on the inlet lip 126 and, if unaddressed, may cause undesirable issues. As such, an anti-icing system 150 may be incorporated into the aircraft engine assembly 100 to prevent formation of ice or to remove formed ice (collectively referenced below as “anti-icing”) from the inlet lip 126.
As described in greater detail below, the anti-icing system 150 circulates heated air in a manner that increases the temperature of portions of the inlet lip 126 that may otherwise be susceptible to icing. The heated air may be provided to the anti-icing system 150 from the engine 104 via a conduit 152, typically as bleed air from a particular stage of the compressor. In one exemplary embodiment, operation of the anti-icing system 150 may be controlled by a controller 154 based on weather conditions, operating conditions, and/or operator commands as a stand-alone system or as part of a larger aircraft safety system. Although not shown, other components, such as flow valves and conduits, may also be incorporated into the aircraft engine assembly 100 to facilitate efficient and effective operation of the anti-icing system 150. The structure and operation of the anti-icing system 150 will be described below in greater detail.
Initially referring to
The anti-icing system 150 may be considered to include an aft (or internal) wall structure 220 that at least partially defines the anti-icing chamber (or inlet lip skin) 210 within the inlet cowl section 120. As shown, the aft wall structure 220 extends in a generally radial direction between the edges of the curved inlet lip 126, although in other embodiments, the aft wall structure 220 may extend between the inner and outer sections 122, 124 or between the one end of the inlet lip 126 and the opposing inner or outer section 122, 124. The inlet lip 126 and wall structure 220 (and the resulting chamber 210) may be arranged and dimensioned with consideration for a number of factors, including the anti-icing function discussed below, as well as engine operation and performance characteristics. As examples, the highlight, or leading edge diameter, of the inlet lip 126 is sized to allow the appropriate amount of air into the engine 104 given the fan diameter size. The length of the inlet lip 126 and engine cowl section 120 may be sized for a certain amount of acoustic treatment. The shape of the inlet lip 126 may also be a function of aerodynamic and manufacturing considerations. Generally, the surface areas of the inlet lip 126 that may be protected from icing are based on the operational characteristics of the aircraft and the engine 104. Although
Considering the cross-sectional shape, the chamber 210 may be referred to as a D-shaped chamber or D-shaped duct.
Again referring to
The nozzle openings 234 may have any suitable orientation. As shown in the depicted exemplary embodiment, the nozzle openings 234 may be positioned along the axial length of the nozzle 230 and clocked or offset with respect to the tangential and/or radial orientations such that the heated air is subject to the desired flow characteristics within the chamber 210, as discussed in greater detail below. In one exemplary embodiment, the nozzle openings 234 are positioned on one tangential side of the nozzle 230, either entirely tangential or with an outwardly radial component to encourage a swirling annular air flow in a designated direction around the chamber 210. In other words, an entirely tangential orientation would be considered to have an angle of 0° relative to a tangential axis. In some embodiments, one or more of the openings 234 may be offset relative to the tangential axis with a radial component. In one exemplary embodiment, a combination of openings 234 with various orientations may be used, including an embodiment with a first opening 234 of 0° relative to the tangential axis, a second opening 234 of 10°-30° relative to the tangential axis, and a third opening 234 of more than 30° relative to the tangential axis. In further embodiments, the openings 234 may have angles of greater than 90°, including 180° or greater. Additionally, the openings 234 may have different sizes.
The anti-icing system 150 further includes a shield 250 with a first end 252 coupled to the aft wall structure 220 and a second end 254 extending into the chamber 210. Like the aft wall structure 220 and the inlet lip 126, the shield 250 has an annular structure extending about the engine assembly centerline 102, as best shown in
The shield 250 functions to divide the chamber 210 into a first chamber portion 212 and a second chamber portion 214. The first chamber portion 212 may be considered to include the portions of the chamber 210 between an inner surface 256 of the shield 250 and the aft wall structure 220, and the second chamber portion 214 may be considered to include the portions of the chamber 210 between an outer surface 258 of the shield 250 and the inlet lip 126. Generally, the first chamber portion 212 may be considered relative to the second chamber portion 214 to be positioned rearward in an axial dimension and inward in a radial dimension. The chamber portions 212, 214 are fluidly coupled together by an annular passage or slot 260 extending between the second end 254 of the shield 250 and the inlet lip 126.
The shield 250 and nozzle 230 are positioned such that the heated air exits the nozzle openings 234 into the first chamber portion 212. The shield 250 is positioned and shaped to at least partially surround and/or enclose the nozzle 230.
The shield 250 may have any suitable shape, position, and dimension to improve the flow characteristics and the overall effectiveness and efficiency of the anti-icing system 150. In one exemplary embodiment, the cross-sectional area (and volume) of the first chamber portion 212 is approximately equal to the cross-sectional area (and volume) of the second chamber portion 214. As further examples, distances 280, 282, 284 may be sized to provide the desired flow characteristics. Generally, the distances 280, 282, 284 and other dimensions are typically defined with CFD analysis. The relative geometric dimensions (e.g., passage 260 to chamber portion 212 size, volume ratio of chamber portion 212 to chamber portion 214, ratio of the swirl nozzle exit area to the chamber portion 212 cross-section area, and the like.). More specific dimensions may depend on the specific engine and application.
As noted above, the shield 250 generally has a curved configuration to partially enclose the nozzle 230, particularly the nozzle openings 234, thereby shielding the inlet lip 126 from direct impingement of the heated air exiting the nozzle openings 234. As a result, the characteristics of the shield 250 may be a function of the characteristics of the nozzle 230, such as the length and position of the nozzle openings 234, the orientation and size of the nozzle openings 234, and the pressure, temperature, and flow rates of the air through the nozzle openings 234.
In the view of
In one exemplary embodiment, the end 252 of the shield 250 is positioned approximately half way or within the inner third along the length of the aft wall 220, although other configurations may be provided. In some embodiments, the axial length of the shield 250 is at least half of the axial length of the chamber 210. In one example, the distances 282, 284 are less than the distance 280 between the nozzle 230 and the inner surface of the inlet lip 126.
As best shown by
The anti-icing system 150 may be formed by any suitable material. In one exemplary embodiment, the aft wall structure 220 and shield 250 may be formed by a material suitable for higher temperatures, such as titanium, while the inlet lip 126 may be formed by a lighter or cheaper material, such as aluminum, that does not need to be as robust relative to elevated temperatures.
An exemplary resulting flow pattern is depicted by
Accordingly, exemplary embodiments discussed above provide improved anti-icing systems and methods. As an example, exemplary embodiments provide a more controlled and efficient use of bleed air for the anti-icing function, especially as compared to conventional systems that merely dump heated air into interior portions of the nacelle. As a further example, exemplary embodiments provide more uniform and effective heating about the circumference of the inlet lip, especially as compared to conventional systems with pipes having numerous piccolo holes, which may result in undesirable hotspots. The swirling rotational flow characteristics provided by the shield enable improved mixing, circulation, and distribution of the injected hot air within the circular leading edge of an engine to prevent ice build-up thereon. Further, such improvements may be achieved with a reduction in cost, complexity, and weight as compared to conventional approaches.
While at least one exemplary embodiment has been presented in the foregoing detailed description of the invention, it should be appreciated that a vast number of variations exist. It should also be appreciated that the exemplary embodiment or exemplary embodiments are only examples, and are not intended to limit the scope, applicability, or configuration of the invention in any way. Rather, the foregoing detailed description will provide those skilled in the art with a convenient road map for implementing an exemplary embodiment of the invention. It being understood that various changes may be made in the function and arrangement of elements described in an exemplary embodiment without departing from the scope of the invention as set forth in the appended claims.
Claims
1. An anti-icing system for an inlet lip annularly extending about a nacelle of an aircraft engine assembly, comprising:
- an interior wall structure at least partially forming an annular anti-icing chamber with an internal surface of the inlet lip;
- an annular shield with a first end coupled to the interior wall structure and a second end extending into the annular anti-icing chamber, the annular shield dividing the annular anti-icing chamber into a first chamber portion and a second chamber portion, the first and second chamber portions being fluidly coupled together by a passage formed between the second end and the internal surface of the inlet lip; and
- a nozzle configured to be coupled to a source of heated air and extending through the interior wall at an inwardly radial position relative to the first end of the annular shield such that the heated air exits the nozzle into the first chamber portion in which the annular shield blocks direct impingement by the heated air on the internal surface of the inlet lip.
2. The anti-icing system of claim 1, wherein the annular shield has an inner surface and an outer surface such that the first chamber portion is formed by the inner surface of the annular shield and the interior wall structure and the second chamber portion is formed by the outer surface of the annular shield and the internal surface of the inlet lip.
3. The anti-icing system of claim 1, wherein the annular anti-icing chamber is a D-shaped chamber.
4. The anti-icing system of claim 1, wherein the nozzle includes a plurality of nozzle openings, including at least a first nozzle opening oriented in a tangential direction.
5. The anti-icing system of claim 4, wherein the plurality of nozzle openings includes a second nozzle opening that is oriented at an angle offset relative to the tangential direction.
6. The anti-icing system of claim 1, wherein the nozzle includes a plurality of nozzle openings, including a first nozzle opening oriented at a first angle of less than 10° relative to a tangential direction, a second nozzle opening oriented at a second angle of between 10° and 30° relative to the tangential direction, and a third nozzle opening oriented at a third angle of greater than 30° relative to the tangential direction.
7. The anti-icing system of claim 1, wherein the annular shield is C-shaped with the second end curving around an end of the nozzle to at least partially enclose the nozzle.
8. The anti-icing system of claim 1, wherein the first chamber portion has a first cross-sectional area and the second chamber portion has a second cross-sectional area, the first cross-sectional area being approximately equal to the second cross-sectional area.
9. The anti-icing system of claim 1, further comprising a plurality of exhaust holes formed about a circumference of the inlet lip.
10. The anti-icing system of claim 1, wherein the interior wall structure and the annular shield are formed from a first material, different from a second material forming the inlet lip.
11. The anti-icing system of claim 1, wherein the annular shield is arranged relative to the nozzle and the interior wall such that the heated air exiting the nozzle completes, on average, at least one revolution around the first chamber portion prior to flowing into the second chamber portion.
12. An engine assembly, comprising:
- an engine;
- a nacelle housing the engine and having a leading portion formed by an inlet lip; and
- an anti-icing system configured to direct air from the engine to the inlet lip for anti-icing the inlet lip, the anti-icing system comprising an interior wall structure at least partially forming an annular anti-icing chamber with an internal surface of the inlet lip; an annular shield with a first end coupled to the interior wall structure and a second end extending into the annular anti-icing chamber, the annular shield dividing the annular anti-icing chamber into a first chamber portion and a second chamber portion, the first and second chamber portions being fluidly coupled together by a passage formed between the second end and the internal surface of the inlet lip; and a nozzle configured to be coupled to a source of heated air and extending through the interior wall at an inwardly radial position relative to the first end of the annular shield such that the heated air exits the nozzle into the first chamber portion in which the annular shield blocks direct impingement by the heated air on the internal surface of the inlet lip.
13. The engine assembly of claim 12, wherein the annular shield has an inner surface and an outer surface such that the first chamber portion is formed by the inner surface of the annular shield and the interior wall structure and the second chamber portion is formed by the outer surface of the annular shield and the internal surface of the inlet lip.
14. The engine assembly of claim 12, wherein the annular anti-icing chamber is a D-shaped chamber.
15. The engine assembly of claim 12, wherein the nozzle includes a plurality of nozzle openings, including a first nozzle opening oriented at a first angle of less than 10° relative to a tangential direction, a second nozzle opening oriented at a second angle of between 10° and 30° relative to the tangential direction, and a third nozzle opening oriented at a third angle of greater than 30° relative to the tangential direction.
16. The engine assembly of claim 11, wherein the annular shield is C-shaped with the second end curving around an end of the nozzle to at least partially enclose the nozzle.
17. The engine assembly of claim 12, wherein the first chamber portion has a first cross-sectional area and the second chamber portion has a second cross-sectional area, the first cross-sectional area being approximately equal to the second cross-sectional area.
18. The engine assembly of claim 12, further comprising a plurality of exhaust holes formed about a circumference of the inlet lip.
19. The engine assembly of claim 12, wherein the interior wall structure and the annular shield are formed from a first material, different from a second material forming the inlet lip.
20. The engine assembly of claim 12, wherein the annular shield is arranged relative to the nozzle and the interior wall such that the heated air exiting the nozzle completes, on average, at least one revolution, on average, around the first chamber portion prior to flowing into the second chamber portion.
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Type: Grant
Filed: Mar 9, 2016
Date of Patent: Apr 24, 2018
Patent Publication Number: 20170259926
Assignee: HONEYWELL INTERNATIONAL INC. (Morris Plains, NJ)
Inventors: Morris Anderson (Mesa, AZ), Shawn Alstad (Peoria, AZ)
Primary Examiner: Valentina Xavier
Application Number: 15/064,748